The present invention relates to a filament-wound combustion chamber for a solid propellant rocket motor, this chamber comprising at its aft opening connection means intended to allow direct mounting of a hot-throughout nozzle made from heat-conducting refractory composite materials.
It is now currently known to produce the combustion chamber of a solid propellant rocket motor by winding glass, Kevlar or carbon fibers, coated with a thermosetting resin, generally an epoxy resin and protected from the heat, on the inside, by a heat-protecting layer in elastomer. In order to establish the connection of the nozzle to the combustion chamber, a metal polar boss (or adapter) is typically used.
The nozzle is equipped with a metallic flange permitting the connection with the metal polar boss of the combustion chamber by means of screws, of a encircling ring or any other adequate means.
To reduce the overall weight, the present trend is to produce hot-throughout nozzles in one or a plurality of components made from a heat-conducting refractory materials of the carbon-carbon type. However, such nozzles need to be insulated against heat at the level of the metallic flange, because the wound portion of the combustion chamber can only withstand very limited heating, about 100.degree. C. This implies a relatively complex design of the nozzle. According to a conventional construction (such as the nozzle of the European MAGE 2 Apogee Motor, for example), the connection of the nozzle with the metal polar boss of the combustion chamber is made by means of a titanium flange surrounded by insulative carbon/phenolic parts (reference numerals 9 and 10 on FIG. 1), blocked on the nozzle by means of a screwed check nut made from a heat-conducting refractory carbon-carbon material.
This assembly of components shows a certain complexity resulting in high production and assembly costs. In addition, the weight of the various components, constituting the combustion chamber and the nozzle, and in particular the metal components, i.e. the aft polar boss of the combustion chamber, the nozzle flange and attaching means (screw, ring, etc.), restricts the possibility of reducing the weight of the rocket motor. Moreover, these metallic components make the rocket motor more likely to be detected by radar means.
The above-mentioned drawbacks, the constraints related to the forging of the metal parts, and their high cost price, have led to searching for new solutions particularly intended to lighten the combustion chamber-nozzle connection, these attempts not meeting with much success up to the present invention. The "Hercules" process described in French Patent No. 2251756 (corresponding to U.S. patent application Ser. No. 416,552 of Nov. 16, 1973) is particularly noted on this point. Said process consists in incorporating fibrous reinforcements in the wound portion of the combustion chamber, in order to be able to connect the nozzle directly to this chamber. Moreover, this process solves the problem only partially, since the nozzle still has to be thermally insulated from the combustion chamber.